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US8353673B2 - Main rotor blade with integral cuff - Google Patents

Main rotor blade with integral cuff Download PDF

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Publication number
US8353673B2
US8353673B2 US12/110,312 US11031208A US8353673B2 US 8353673 B2 US8353673 B2 US 8353673B2 US 11031208 A US11031208 A US 11031208A US 8353673 B2 US8353673 B2 US 8353673B2
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US
United States
Prior art keywords
main
spar
tip
section
recited
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Expired - Fee Related, expires
Application number
US12/110,312
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English (en)
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US20090269205A1 (en
Inventor
Kevin P. Leahy
Richard Joseph Simkulak
David A. Kovalsky
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Sikorsky Aircraft Corp
Original Assignee
Sikorsky Aircraft Corp
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Sikorsky Aircraft Corp filed Critical Sikorsky Aircraft Corp
Priority to US12/110,312 priority Critical patent/US8353673B2/en
Assigned to SIKORSKY AIRCRAFT CORPORATION reassignment SIKORSKY AIRCRAFT CORPORATION ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: KOVALSKY, DAVID A., LEAHY, KEVIN P., SIMKULAK, RICHARD JOSEPH
Priority to EP08874007.1A priority patent/EP2276666B1/fr
Priority to PCT/US2008/084455 priority patent/WO2009131600A2/fr
Publication of US20090269205A1 publication Critical patent/US20090269205A1/en
Application granted granted Critical
Publication of US8353673B2 publication Critical patent/US8353673B2/en
Expired - Fee Related legal-status Critical Current
Adjusted expiration legal-status Critical

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Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C27/00—Rotorcraft; Rotors peculiar thereto
    • B64C27/32—Rotors
    • B64C27/46—Blades
    • B64C27/463—Blade tips
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C27/00—Rotorcraft; Rotors peculiar thereto
    • B64C27/32—Rotors
    • B64C27/46—Blades
    • B64C27/473—Constructional features
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C27/00—Rotorcraft; Rotors peculiar thereto
    • B64C27/32—Rotors
    • B64C27/46—Blades
    • B64C27/473—Constructional features
    • B64C2027/4733—Rotor blades substantially made from particular materials
    • B64C2027/4736—Rotor blades substantially made from particular materials from composite materials

Definitions

  • the present invention relates to a rotary wing aircraft main rotor blade, and more particularly to a main rotor blade with a cuff integral with the spar.
  • Rotary-wing aircraft rotor blades often include a composite or metallic main rotor spar of a constant chord and constant wall thickness.
  • a separate root end cuff is attached to an inboard section of the main rotor spar.
  • the cuff is conventionally manufactured of a titanium forging or a composite root laminate sized to match an outer mold line of the main rotor spar.
  • the root end cuff is attached to the main rotor spar through a multiple fastener configuration, each fastener of which must be torqued to a required standard.
  • the separate root end cuff assembly and related hub componentry may result in a weight and life cycle penalty as the multiple of components are duplicated on each rotor blade assembly.
  • a main rotor blade assembly includes: a main spar formed with an integral cuff.
  • a rotary wing aircraft includes: a rotor hub assembly; a main spar which comprises a root section and an main section, the root section forms an integral cuff, the integral cuff having a first hub mount aperture and a second hub mount aperture; a first pin which passes through the first aperture to mount the main spar to the rotor hub assembly; and a second pin which passes through the second aperture to mount the main spar to the rotor hub assembly.
  • a method of mounting a main rotor blade assembly includes: attaching an integral cuff directly to a rotor hub assembly.
  • FIG. 1 is a perspective view of a rotary wing aircraft for use with the present invention
  • FIG. 2A is a perspective view of a main rotor blade
  • FIG. 2B is an expanded perspective view of a tip section of the rotor blade of FIG. 2A ;
  • FIG. 2C is an expanded rear oblique perspective view of a tip section of the rotor blade of FIG. 2A ;
  • FIG. 2D is an expanded front oblique perspective view of a tip section of the rotor blade of FIG. 2A ;
  • FIG. 2E is an expanded top view of a tip section of the rotor blade of FIG. 2A ;
  • FIG. 2F is an expanded front view of a tip section of the rotor blade of FIG. 2A ;
  • FIG. 3 is an exploded view of a main rotor blade
  • FIG. 4 is an exploded view of a removable tip cap for a main rotor blade
  • FIG. 5A is a plan view of a tip section of a main rotor blade
  • FIG. 5B is a top plan view of a main rotor blade
  • FIG. 5C is an exploded view of a tip section of the main rotor blade
  • FIG. 5D is an expanded view of a distal end of a tip spar assembly without a tip cap mounted thereto;
  • FIG. 5E is an expanded section view of a tip spar bonded to a main blade spar with a tip cap installed to the tip spar;
  • FIG. 5F is an exploded perspective view of a main spar with a forward tip spar section and an aft tip spar section;
  • FIG. 5G is an assembled view of the main spar, forward tip spar section and aft tip spar section;
  • FIG. 5H is an exploded view of an assembled main spar and tip spar with a main core and tip core;
  • FIG. 5I is a perspective view of an assembled main spar and tip spar with the main core and tip core;
  • FIG. 5J is an exploded view of the main rotor blade assembly components which include a non-straight tip form
  • FIG. 5K is a front exploded view of the main rotor blade assembly components which include a non-straight tip form
  • FIG. 6 is a chart illustrating one manufacturing plan of a main rotor blade in accordance with one non-limiting embodiment of the present invention.
  • FIG. 7A is a top view of the main rotor blade spar
  • FIG. 7B is a front view of the main rotor blade spar
  • FIG. 8A is a sectional view through a root section of the main rotor blade spar taken along line 8 A- 8 A in FIG. 7A ;
  • FIG. 8B is a sectional view through a main section of the main rotor blade spar taken along line 8 B- 8 B in FIG. 7A ;
  • FIG. 8C is a sectional view through an outboard section of the main rotor blade spar taken along line 8 C- 8 C in FIG. 7A ;
  • FIG. 8D is a sectional view through a tip section of the main rotor blade spar taken along line 8 D- 8 D in FIG. 7A ;
  • FIG. 9A is a longitudinal sectional view through the main rotor blade spar
  • FIG. 9B is an expanded sectional view of an inboard section of the main rotor blade spar of FIG. 9A ;
  • FIG. 10A is a top view of the main rotor blade spar
  • FIG. 10B is an expanded top view of an outboard section of the main rotor blade spar of FIG. 10A ;
  • FIG. 11 is a top view of the main rotor blade assembly
  • FIG. 12A is a top view of a main rotor system for a rotary wing aircraft.
  • FIG. 12B is an expanded top view of the main rotor system illustrating an integral cuff of the main rotor blade spar.
  • FIG. 1 schematically illustrates a rotary-wing aircraft 10 having a main rotor system 12 .
  • the aircraft 10 includes an airframe 14 having an extending tail 16 which mounts a tail rotor system 18 , such as an anti-torque system.
  • the main rotor assembly 12 is driven about an axis of rotation A through a main gearbox (illustrated schematically at T) by one or more engines E.
  • the main rotor system 12 includes a multiple of rotor blade assemblies 20 mounted to a rotor hub assembly H.
  • helicopter configuration is illustrated and described in the disclosed non-limiting embodiment, other configurations and/or machines, such as high speed compound rotary wing aircraft with supplemental translational thrust systems, dual contra-rotating, coaxial rotor system aircraft, turbo-props, tilt-rotors and tilt-wing aircraft, will also benefit from the present invention.
  • each rotor blade assembly 20 of the rotor assembly 12 generally includes a root section 22 , an intermediate section 24 , a tip section 26 and a tip cap 28 .
  • Each rotor blade section 22 , 24 , 26 , 28 may define particular airfoil geometries to particularly tailor the rotor blade aerodynamics to the velocity increase along the rotor blade span.
  • the rotor blade tip section 26 includes an anhedral form 27 ( FIG. 2B-2F ), however, any angled and non-angled forms such as cathedral, gull, bent, and other non-straight forms will benefit from the present invention.
  • the anhedral form 27 as defined herein may include a rotor blade tip section 26 which is at least partially not contained in a plane defined by the intermediate section 24 .
  • the rotor blade sections 22 - 28 define a span R of the main rotor blade assembly 20 between the axis of rotation A and a distal end 30 of the tip cap 28 such that any radial station may be expressed as a percentage in terms of a blade radius x/R.
  • the rotor blade assembly 20 defines a longitudinal feathering axis P between a leading edge 32 and a trailing edge 34 .
  • the distance between the leading edge 32 and the trailing edge 34 defines a main element chord length Cm.
  • the rotor blade assembly 20 generally includes a main blade assembly 40 and a tip assembly 42 .
  • the main blade assembly 40 includes an upper skin 44 a main core 46 , a main spar 48 , a lower skin 50 , and a leading edge assembly 52 .
  • the main spar 48 , main core 46 and skins 44 , 50 are generally referred to as a pocket assembly, the forward portion of which is closed out by the leading edge assembly 52 .
  • the leading edge assembly 52 generally includes a main sheath laminate 54 upon which is mounted a wear-resistant material such as a titanium erosion strip 56 and a nickel erosion strip 58 to provide abrasion protection.
  • the sheath laminate 54 may include a single erosion strip 60 manufactured of AM355 which replaces the titanium erosion strip 56 and the nickel erosion strip 58 .
  • a wear-resistant materials may alternatively or additionally be provided for the leading edge assembly 52 .
  • Additional structures such as weight cups 62 , blade tie down fastener structures 64 and trim tab systems 66 may also be provided, further details of which need not be further elaborated herein.
  • the tip assembly 42 generally includes a tip spar 48 T, a tip core 46 T, a tip leading edge assembly 68 and the tip cap 28 .
  • the tip cap 28 may be removably attached to the tip spar 48 T though hardware 70 in a cap interface 78 ( FIG. 4 ) to close-out the end of the tip assembly 42 .
  • the cap interface 78 may be defined at an inboard radial station 77 of approximately ninety eight percent (98% x/R) and end at an outboard radial station 79 which defines the distal end of the tip spar 48 T, however other locations may alternatively be provided. It should be understood that other overlap lengths, locations and interfaces may alternatively or additionally be provided.
  • the tip leading edge assembly 68 may also include a splice cap laminate 54 T and a tip erosion strip 72 to provide abrasion protection.
  • the splice cap laminate 54 T abuts the main sheath laminate 54 at the inboard radial station 82 .
  • the tip erosion strip 72 may overlap the inboard radial station 82 and extend to the outboard radial station 84 to protect the interface 78 .
  • the tip cap 28 also includes an erosion cap 76 which may overlap the tip cap interface 77 between the tip assembly 42 and the tip cap 28 .
  • the tip erosion strip 72 and the erosion cap 76 may include AM355, titanium, nickel, or a variety of other wear-resistant materials or combinations thereof.
  • the main spar 48 and the tip spar 48 T extend along the length of the blade sections 22 - 26 .
  • the main spar 48 and the tip spar 48 T are structural members manufactured of a high strength non-metallic composite material and/or high strength metallic material such as titanium.
  • the tip spar 48 T is attached to the main spar 48 in a male-female overlap relationship 80 between the inboard radial station 82 and the outboard radial station 84 .
  • the upper skin 44 and the lower skin 50 extend outboard of the overlap relationship 80 which allows a contiguous composite skin laminate from the root section 22 through the tip section 26 to define the aerodynamic shape of the blade ( FIG. 5B ). That is, no separate main skins and tip skins are utilized.
  • the upper skin 44 and the lower skin 50 extend from the root section 22 to the inboard radial station 77 of the cap interface 78 of the tip spar 48 T which abuts the tip cap 28 .
  • the upper skin 44 and the lower skin 50 extend past the overlap relationship 80 .
  • Lift properties are significantly increased over conventional rotor blades with separate tip skins due in part to the increased strength of the one-piece composite skin laminate. Should a blade strike occur, the tip assembly 42 may still separate from the main blade assembly 40 at the overlap relationship 80 such that only the tip spar 48 T and related tip assembly 42 components may break away which thereby allows the main blade 20 defined by the main spar 48 to remain flightworthy, albeit with reduced lift capabilities.
  • the tip spar 48 T and the tip core 46 T define the anhedral form or other angled forms such as cathedral, gull, bent, and others. Furthermore, the tip core 46 T may be hot formed to specifically include the anhedral form as well as be of similar or different weight per cubic foot (pcf) as compared to the main core 46 to provide particular lift and strength properties.
  • the anhedral form may be located at a radial station of approximately ninety six percent (96% x/R). It should be understood that the anhedral form or other angled and non-angled forms may be located at other radial stations.
  • An end section ST of the tip spar 48 T is bonded over an end section SM of the main spar 48 ( FIG. 5C ) to define the overlap relationship 80 .
  • An adhesive material such as epoxy film adhesive, is one example of a bonding agent.
  • the main spar end section SM may be of a reduced perimeter such that a flush outer mold line (OML) spar surface is formed therebetween. That is, the end section SM of the main spar 48 defines a jogged or stepped area to receive the end section ST of the tip spar 48 T.
  • the overlap relationship 80 may begin at an inboard radial station 82 of approximately eighty nine percent (89% x/R) and end at an outboard radial station 84 of approximately ninety two percent (92% x/R). It should be understood that other overlap length, locations and interfaces may alternatively or additionally be provided.
  • the tip spar 48 T may include a forward tip spar section 48 Tf and an aft tip spar section 48 Ta.
  • the forward tip spar section 48 Tf includes an open aft section 90 with a stepped inner mold line to receive the open aft section 90 of the forward tip spar section 48 Tf of FIG. 5D .
  • the aft tip spar section 48 Ta includes an open forward section 92 with a stepped outer mold line. It should be understood that other overlap length, locations and interfaces may alternatively or additionally be provided.
  • the forward and aft split arrangement facilitates assembly of the forward tip spar section 48 Tf to the aft tip spar section 48 Ta such that the inboard end section ST of the tip spar 48 T is receivable over the outboard end section SM of the main spar 48 ( FIG. 5E ).
  • the forward tip spar section 48 Tf and an aft tip spar section 48 Ta are thereby assembled together over the outboard end section SM of the main spar 48 ( FIG. 5F ) to attach the tip spar 48 T to the main spar 48 ( FIG. 5G ).
  • a core interface 46 I between the main core 46 and the tip core 46 T is offset inboard from the inboard radial station 82 of the overlap relationship 80 .
  • the main core 46 and the tip core 46 T include a respective leading edge contour 46 c and 46 Tc which generally match a trailing edge 48 A of the main spar 48 and a trailing edge 48 TA of the tip spar 48 T to provide an overlapped interface ( FIG. 5I ) which define an outer contour to receive the upper skin 44 and the lower skin 50 ( FIG. 5J ).
  • the rotor blade anhedral is integrated into the upper skin 44 , the tip core 46 T, the tip spar 48 T and the lower skin 50 ( FIG. 5K ). That is, the angled form is directly formed into the upper skin 44 , the tip core 46 T, the tip spar 48 T, and the lower skin 50 to provide the tip section 26 with increased strength and lift properties. Sweep is also integrated into the tip section 26 .
  • Assembly and manufacture of the main rotor blade assembly 20 may generally follow the manufacturing plan as schematically illustrated in FIG. 6 .
  • the main spar 48 generally includes a root section 100 , a main section 102 , an outboard section 104 and a tip section 106 .
  • the root section 100 is formed as an integral cuff 108 which defines an increased root wall thickness ( FIG. 8A ) relative a main wall thickness ( FIG. 8B ) within the main section 102 , the outboard wall thickness ( FIG. 8C ) within the outboard section 104 and a tip wall thickness ( FIG. 8D ) within the tip section 106 .
  • the root wall thickness is twice as thick as the main wall thickness (also illustrated in FIGS. 9A and 9B ).
  • the root wall thickness tapers into the main wall thickness as the root section 100 tapers chord-wise and thickness-wise.
  • the main section 102 ( FIG. 8B ) shape and chord-wise dimensions are generally equivalent to the outboard section 104 ( FIG. 8C ) which has a slightly decreased thickness with a transition area taper section 81 therebetween.
  • the transition area taper section 81 provides for the airfoil shape in the main section 102 to the outboard section 104 .
  • the tip wall thickness ( FIG. 8D ) within the tip section 106 tapers at a taper section 83 relative to the outboard section 104 (also illustrated in FIGS.
  • the integral cuff 108 of the root section 100 of each rotor blade assembly 20 includes a first aperture 110 A and a second aperture 110 B such that each rotor blade assembly 20 is mounted directly to the rotor hub assembly H ( FIG. 12A ).
  • the first aperture 110 A and the second aperture 110 B receive a first pin 112 A and a second pin 112 B to directly attach an integral hinge cuff 114 to the main spar 48 . That is, the root section 100 forms the integral cuff 108 for direct mount of the rotor blade assembly 20 without the conventional root laminate and integral cuff plate which are bonded to the spar. That is, conventional cuff attachment assemblies are eliminated.
  • the integral hinge cuff 114 is mounted to a sleeve 116 which is then mounted to a rotor hub 118 through a yoke 120 .

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
US12/110,312 2008-04-26 2008-04-26 Main rotor blade with integral cuff Expired - Fee Related US8353673B2 (en)

Priority Applications (3)

Application Number Priority Date Filing Date Title
US12/110,312 US8353673B2 (en) 2008-04-26 2008-04-26 Main rotor blade with integral cuff
EP08874007.1A EP2276666B1 (fr) 2008-04-26 2008-11-23 Pâle de rotor principal à manchon intégral
PCT/US2008/084455 WO2009131600A2 (fr) 2008-04-26 2008-11-23 Pâle de rotor principal à manchon intégral

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Application Number Priority Date Filing Date Title
US12/110,312 US8353673B2 (en) 2008-04-26 2008-04-26 Main rotor blade with integral cuff

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US20090269205A1 US20090269205A1 (en) 2009-10-29
US8353673B2 true US8353673B2 (en) 2013-01-15

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EP (1) EP2276666B1 (fr)
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US9745056B2 (en) 2013-05-14 2017-08-29 Sikorsky Aircraft Corporation Main rotor blade with composite integral skin and cuff
US9981736B2 (en) 2016-05-16 2018-05-29 Bell Helicopter Textron Inc. Folding proprotor blade assembly having integral tang assembly
US10239604B2 (en) * 2016-05-21 2019-03-26 Bell Helicopter Textron Inc. Structurally biased proprotor blade assembly
US10301013B2 (en) 2016-08-09 2019-05-28 Sikorsky Aircraft Corporation Bond fixture for composite splice cap assembly
US10351229B2 (en) 2014-05-15 2019-07-16 Sikorsky Aircraft Corporation Metallic dimpled doubler
US20200277085A1 (en) * 2016-08-09 2020-09-03 Sikorsky Aircraft Corporation Main rotor blade cuff bond fixture
US20210094682A1 (en) * 2016-09-30 2021-04-01 Sikorsky Aircraft Corporation De-ice fairing bond fixture

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US20110243750A1 (en) 2010-01-14 2011-10-06 Neptco, Inc. Wind Turbine Rotor Blade Components and Methods of Making Same
US10137542B2 (en) 2010-01-14 2018-11-27 Senvion Gmbh Wind turbine rotor blade components and machine for making same
US8858183B2 (en) * 2010-02-26 2014-10-14 Sikorsky Aircraft Corporation Rotor blade for a rotary-wing aircraft
US9429025B2 (en) * 2010-06-22 2016-08-30 Sikorsky Aircraft Corporation Erosion resistant helicopter blade
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US10689104B2 (en) * 2017-05-09 2020-06-23 Textron Innovations Inc. Tail rotor integrated damper attachment
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US20240003329A1 (en) * 2022-07-01 2024-01-04 Rohr, Inc. Composite thermoplastic rotor blade with integral cuff
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Cited By (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US9745056B2 (en) 2013-05-14 2017-08-29 Sikorsky Aircraft Corporation Main rotor blade with composite integral skin and cuff
US10351229B2 (en) 2014-05-15 2019-07-16 Sikorsky Aircraft Corporation Metallic dimpled doubler
US9981736B2 (en) 2016-05-16 2018-05-29 Bell Helicopter Textron Inc. Folding proprotor blade assembly having integral tang assembly
US10239604B2 (en) * 2016-05-21 2019-03-26 Bell Helicopter Textron Inc. Structurally biased proprotor blade assembly
US10301013B2 (en) 2016-08-09 2019-05-28 Sikorsky Aircraft Corporation Bond fixture for composite splice cap assembly
US20200277085A1 (en) * 2016-08-09 2020-09-03 Sikorsky Aircraft Corporation Main rotor blade cuff bond fixture
US11220331B2 (en) 2016-08-09 2022-01-11 Sikorsky Aircraft Corporation Bond fixture for composite splice cap assembly
US11834200B2 (en) * 2016-08-09 2023-12-05 Sikorsky Aircraft Corporation Method of bonding a blade cuff to a rotor blade
US20210094682A1 (en) * 2016-09-30 2021-04-01 Sikorsky Aircraft Corporation De-ice fairing bond fixture
US12071229B2 (en) * 2016-09-30 2024-08-27 Sikorsky Aircraft Corporation De-ice fairing bond fixture

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US20090269205A1 (en) 2009-10-29
EP2276666A2 (fr) 2011-01-26
EP2276666A4 (fr) 2013-08-28
EP2276666B1 (fr) 2017-10-11
WO2009131600A2 (fr) 2009-10-29
WO2009131600A3 (fr) 2010-01-07

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